Gas turbine engine component with diffusive cooling hole

ABSTRACT

A component for a gas turbine engine includes a gas path wall having a first surface, a second surface exposed to hot gas flow, and a cooling hole extending through the gas path wall. The cooling hole includes an inlet formed in the first surface, an outlet formed in the second surface, cooling hole surfaces that define the cooling hole between the inlet and the outlet, and a longitudinal ridge formed along at least one of the cooling hole surfaces. The longitudinal ridge separates the cooling hole into first and second lobes. The cooling hole diverges through the gas path wall, such that cross-sectional area of the cooling hole increases continuously from the inlet through the cooling hole to the outlet.

CROSS REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Application No.61/599,242, filed on Feb. 15, 2012, and entitled “Gas Turbine EngineComponent with Diffusive Hole,” the disclosure of which is incorporatedby reference in its entirety.

BACKGROUND

This invention relates generally to turbomachinery, and specifically toturbine flow path components for gas turbine engines. In particular, theinvention relates to cooling techniques for airfoils and other gasturbine engine components exposed to hot working fluid flow, including,but not limited to, rotor blades and stator vane airfoils, endwallsurfaces including platforms, shrouds and compressor and turbinecasings, combustor liners, turbine exhaust assemblies, thrust augmentorsand exhaust nozzles.

Gas turbine engines are rotary-type combustion turbine engines builtaround a power core made up of a compressor, combustor and turbine,arranged in flow series with an upstream inlet and downstream exhaust.The compressor section compresses air from the inlet, which is mixedwith fuel in the combustor and ignited to generate hot combustion gas.The turbine section extracts energy from the expanding combustion gas,and drives the compressor section via a common shaft. Expandedcombustion products are exhausted downstream, and energy is delivered inthe form of rotational energy in the shaft, reactive thrust from theexhaust, or both.

Gas turbine engines provide efficient, reliable power for a wide rangeof applications in aviation, transportation and industrial powergeneration. Small-scale gas turbine engines typically utilize aone-spool design, with co-rotating compressor and turbine sections.Larger-scale combustion turbines including jet engines and industrialgas turbines (IGTs) are generally arranged into a number of coaxiallynested spools. The spools operate at different pressures, temperaturesand spool speeds, and may rotate in different directions.

Individual compressor and turbine sections in each spool may also besubdivided into a number of stages, formed of alternating rows of rotorblade and stator vane airfoils. The airfoils are shaped to turn,accelerate and compress the working fluid flow, or to generate lift forconversion to rotational energy in the turbine.

Industrial gas turbines often utilize complex nested spoolconfigurations, and deliver power via an output shaft coupled to anelectrical generator or other load, typically using an external gearbox.In combined cycle gas turbines (CCGTs), a steam turbine or othersecondary system is used to extract additional energy from the exhaust,improving thermodynamic efficiency. Gas turbine engines are also used inmarine and land-based applications, including naval vessels, trains andarmored vehicles, and in smaller-scale applications such as auxiliarypower units.

Aviation applications include turbojet, turbofan, turboprop andturboshaft engine designs. In turbojet engines, thrust is generatedprimarily from the exhaust. Modern fixed-wing aircraft generally employturbofan and turboprop configurations, in which the low pressure spoolis coupled to a propulsion fan or propeller. Turboshaft engines areemployed on rotary-wing aircraft, including helicopters, typically usinga reduction gearbox to control blade speed. Unducted (open rotor)turbofans and ducted propeller engines also known, in a variety ofsingle-rotor and contra-rotating designs with both forward and aftmounting configurations.

Aviation turbines generally utilize two and three-spool configurations,with a corresponding number of coaxially rotating turbine and compressorsections. In two-spool designs, the high pressure turbine drives a highpressure compressor, forming the high pressure spool or high spool. Thelow-pressure turbine drives the low spool and fan section, or a shaftfor a rotor or propeller. In three-spool engines, there is also anintermediate pressure spool. Aviation turbines are also used to powerauxiliary devices including electrical generators, hydraulic pumps andelements of the environmental control system, for example using bleedair from the compressor or via an accessory gearbox.

Additional turbine engine applications and turbine engine types includeintercooled, regenerated or recuperated and variable cycle gas turbineengines, and combinations thereof. In particular, these applicationsinclude intercooled turbine engines, for example with a relativelyhigher pressure ratio, regenerated or recuperated gas turbine engines,for example with a relatively lower pressure ratio or for smaller-scaleapplications, and variable cycle gas turbine engines, for example foroperation under a range of flight conditions including subsonic,transonic and supersonic speeds. Combined intercooled andregenerated/recuperated engines are also known, in a variety of spoolconfigurations with traditional and variable cycle modes of operation.

Turbofan engines are commonly divided into high and low bypassconfigurations. High bypass turbofans generate thrust primarily from thefan, which accelerates airflow through a bypass duct oriented around theengine core. This design is common on commercial aircraft andtransports, where noise and fuel efficiency are primary concerns. Thefan rotor may also operate as a first stage compressor, or as apre-compressor stage for the low-pressure compressor or booster module.Variable-area nozzle surfaces can also be deployed to regulate thebypass pressure and improve fan performance, for example during takeoffand landing. Advanced turbofan engines may also utilize a geared fandrive mechanism to provide greater speed control, reducing noise andincreasing engine efficiency, or to increase or decrease specificthrust.

Low bypass turbofans produce proportionally more thrust from the exhaustflow, generating greater specific thrust for use in high-performanceapplications including supersonic jet aircraft. Low bypass turbofanengines may also include variable-area exhaust nozzles and afterburneror augmentor assemblies for flow regulation and short-term thrustenhancement. Specialized high-speed applications include continuouslyafterburning engines and hybrid turbojet/ramjet configurations.

Across these applications, turbine performance depends on the balancebetween higher pressure ratios and core gas path temperatures, whichtend to increase efficiency, and the related effects on service life andreliability due to increased stress and wear. This balance isparticularly relevant to gas turbine engine components in the hotsections of the compressor, combustor, turbine and exhaust sections,where active cooling is required to prevent damage due to high gas pathtemperatures and pressures.

SUMMARY

A component for a gas turbine engine includes a gas path wall having afirst surface, a second surface exposed to hot gas flow, and a coolinghole extending through the gas path wall. The cooling hole includes aninlet formed in the first surface, an outlet formed in the secondsurface, cooling hole surfaces that define the cooling hole between theinlet and the outlet, and a longitudinal ridge formed along at least oneof the cooling hole surfaces. The longitudinal ridge separates thecooling hole into first and second lobes. The cooling hole divergesthrough the gas path wall, such that cross-sectional area of the coolinghole increases continuously from the inlet through the cooling hole tothe outlet.

Another embodiment of the present invention is an airfoil including aflow path wall having a first surface exposed to cooling fluid and asecond surface exposed to hot gas flow. A cooling hole is formed in theflow path wall and is laterally diverging continuously from an inlet atthe first surface to an outlet at the second surface. A longitudinalridge is defined along the cooling hole and divides the cooling holeinto first and second lobes. Flow of the cooling fluid is substantiallydiffusive through the cooling hole, from the inlet at the first surfaceof the flow path wall through to the outlet at the second surface of theflow path wall.

Another embodiment of the present invention is a gas turbine enginecomponent including a gas path wall, a cooling hole extending throughthe gas path wall, and a longitudinal ridge extending along the coolinghole. The gas path wall has a first surface and a second surface exposedto hot gas flow. The cooling hole is continuously diverging from aninlet in the first surface to an outlet in the second surface. Thelongitudinal ridge divides the cooling hole into first and second lobes.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional view of a gas turbine engine.

FIG. 2A is a perspective view of an airfoil for the gas turbine engine,in a rotor blade configuration.

FIG. 2B is a perspective view of an airfoil for the gas turbine engine,in a stator vane configuration.

FIG. 3A is a cross-sectional view of the gas path wall for a cooled gasturbine engine component, taken in a longitudinal direction.

FIG. 3B is a cross-sectional view of the gas path wall, showing atruncated lobe configuration.

FIG. 3C is a cross-sectional view of the gas path wall, showing a cuspedinlet configuration.

FIG. 4A is a schematic view of the gas path wall, with a three-lobecooling hole geometry.

FIG. 4B is a schematic view of the gas path wall, with a two-lobecooling hole geometry.

FIG. 5A is a schematic view of the gas path wall, with a truncated lobecooling hole geometry.

FIG. 5B is a schematic view of the gas path wall, with an alternatetruncated lobe cooling hole geometry.

FIG. 6A is a schematic view of the gas path wall, with a truncatedtwo-lobe cooling hole geometry.

FIG. 6B is a schematic view of the gas path wall, with a cusped inletcooling hole geometry.

FIG. 7A is a schematic view of the gas path wall, with a three-lobecusped inlet cooling hole geometry.

FIG. 7B is a schematic view of the gas path wall, with a buried dividercooling hole geometry.

FIG. 7C is a schematic view of the gas path wall, with a three-lobetwo-cusped inlet cooling hole geometry.

FIG. 8 is a block diagram of a method for forming a cooling hole in agas turbine engine component.

DETAILED DESCRIPTION

FIG. 1 is a cross-sectional view of gas turbine engine 10. Gas turbineengine (or turbine engine) 10 includes a power core with compressorsection 12, combustor 14 and turbine section 16 arranged in flow seriesbetween upstream inlet 18 and downstream exhaust 20. Compressor section12 and turbine section 16 are arranged into a number of alternatingstages of rotor airfoils (or blades) 22 and stator airfoils (or vanes)24.

In the turbofan configuration of FIG. 1, propulsion fan 26 is positionedin bypass duct 28, which is coaxially oriented about the engine corealong centerline (or turbine axis) C_(L). An open-rotor propulsion stage26 may also provided, with turbine engine 10 operating as a turboprop orunducted turbofan engine. Alternatively, fan rotor 26 and bypass duct 28may be absent, with turbine engine 10 configured as a turbojet orturboshaft engine, or an industrial gas turbine.

For improved service life and reliability, components of gas turbineengine 10 are provided with an improved cooling configuration, asdescribed below. Suitable components for the cooling configurationinclude rotor airfoils 22, stator airfoils 24 and other gas turbineengine components exposed to hot gas flow, including, but not limitedto, platforms, shrouds, casings and other endwall surfaces in hotsections of compressor 12 and turbine 16, and liners, nozzles,afterburners, augmentors and other gas wall components in combustor 14and exhaust section 20.

In the two-spool, high bypass configuration of FIG. 1, compressorsection 12 includes low pressure compressor (LPC) 30 and high pressurecompressor (HPC) 32, and turbine section 16 includes high pressureturbine (HPT) 34 and low pressure turbine (LPT) 36. Low pressurecompressor 30 is rotationally coupled to low pressure turbine 36 via lowpressure (LP) shaft 38, forming the LP spool or low spool. High pressurecompressor 32 is rotationally coupled to high pressure turbine 34 viahigh pressure (HP) shaft 40, forming the HP spool or high spool.

Flow F at inlet 18 divides into primary (core) flow F_(P) and secondary(bypass) flow F_(S) downstream of fan rotor 26. Fan rotor 26 acceleratessecondary flow F_(S) through bypass duct 28, with fan exit guide vanes(FEGVs) 42 to reduce swirl and improve thrust performance. In somedesigns, structural guide vanes (SGVs) 42 are used, providing combinedflow turning and load bearing capabilities.

Primary flow F_(P) is compressed in low pressure compressor 30 and highpressure compressor 32, then mixed with fuel in combustor 14 and ignitedto generate hot combustion gas. The combustion gas expands to providerotational energy in high pressure turbine 34 and low pressure turbine36, driving high pressure compressor 32 and low pressure compressor 30,respectively. Expanded combustion gases exit through exhaust section (orexhaust nozzle) 20, which can be shaped or actuated to regulate theexhaust flow and improve thrust performance.

Low pressure shaft 38 and high pressure shaft 40 are mounted coaxiallyabout centerline C_(L), and rotate at different speeds. Fan rotor (orother propulsion stage) 26 is rotationally coupled to low pressure shaft38. In advanced designs, fan drive gear system 44 is provided foradditional fan speed control, improving thrust performance andefficiency with reduced noise output.

Fan rotor 26 may also function as a first-stage compressor for gasturbine engine 10, and LPC 30 may be configured as an intermediatecompressor or booster. Alternatively, propulsion stage 26 has an openrotor design, or is absent, as described above. Gas turbine engine 10thus encompasses a wide range of different shaft, spool and turbineengine configurations, including one, two and three-spool turboprop and(high or low bypass) turbofan engines, turboshaft engines, turbojetengines, and multi-spool industrial gas turbines.

In each of these applications, turbine efficiency and performance dependon the overall pressure ratio, defined by the total pressure at inlet 18as compared to the exit pressure of compressor section 12, for exampleat the outlet of high pressure compressor 32, entering combustor 14.Higher pressure ratios, however, also result in greater gas pathtemperatures, increasing the cooling loads on rotor airfoils 22, statorairfoils 24 and other components of gas turbine engine 10. To reduceoperating temperatures, increase service life and maintain engineefficiency, these components are provided with improved coolingconfigurations, as described below. Suitable components include, but arenot limited to, cooled gas turbine engine components in compressorsections 30 and 32, combustor 14, turbine sections 34 and 36, andexhaust section 20 of gas turbine engine 10.

FIG. 2A is a perspective view of rotor airfoil (or blade) 22 for gasturbine engine 10, as shown in FIG. 1, or for another turbomachine.Rotor airfoil 22 extends axially from leading edge 51 to trailing edge52, defining pressure surface 53 (front) and suction surface 54 (back)therebetween.

Pressure and suction surfaces 53 and 54 form the major opposing surfacesor walls of airfoil 22, extending axially between leading edge 51 andtrailing edge 52, and radially from root section 55, adjacent innerdiameter (ID) platform 56, to tip section 57, opposite ID platform 56.In some designs, tip section 57 is shrouded.

Cooling holes or outlets 60 are provided on one or more surfaces ofairfoil 22, for example along leading edge 51, trailing edge 52,pressure (or concave) surface 53, or suction (or convex) surface 54, ora combination thereof. Cooling holes or passages 60 may also be providedon the endwall surfaces of airfoil 22, for example along ID platform 56,or on a shroud or engine casing adjacent tip section 57.

FIG. 2B is a perspective view of stator airfoil (or vane) 24 for gasturbine engine 10, as shown in FIG. 1, or for another turbomachine.Stator airfoil 24 extends axially from leading edge 61 to trailing edge62, defining pressure surface 63 (front) and suction surface 64 (back)therebetween. Pressure and suction surfaces 63 and 64 extend from inner(or root) section 65, adjacent ID platform 66, to outer (or tip) section67, adjacent outer diameter (OD) platform 68.

Cooling holes or outlets 60 are provided along one or more surfaces ofairfoil 24, for example leading or trailing edge 61 or 62, pressure(concave) or suction (convex) surface 63 or 64, or a combinationthereof. Cooling holes or passages 60 may also be provided on theendwall surfaces of airfoil 24, for example along ID platform 66 and ODplatform 68.

Rotor airfoils 22 (FIG. 2A) and stator airfoils 24 (FIG. 2B) are formedof high strength, heat resistant materials such as high temperaturealloys and superalloys, and are provided with thermal anderosion-resistant coatings. Airfoils 22 and 24 are also provided withinternal cooling holes and cooling holes 60 to reduce thermal fatigueand wear, and to prevent melting when exposed to hot gas flow in thehigher temperature regions of a gas turbine engine or otherturbomachine. Cooling holes 60 deliver cooling fluid (e.g., steam or airfrom a compressor) through the outer walls and platform structures ofairfoils 22 and 24, creating a thin layer (or film) of cooling fluid toprotect the outer (gas path) surfaces from high temperature flow.

While surface cooling extends service life and increases reliability,injecting cooling fluid into the gas path also reduces engineefficiency, and the cost in efficiency increases with the requiredcooling flow. Cooling holes 60 are thus provided with improved meteringand inlet geometry to reduce jets and blow off, and improved diffusionand exit geometry to reduce flow separation and corner effects. Coolingholes 60 reduce flow requirements and improve the spread of coolingfluid across the hot surfaces of airfoils 22 and 24, and other gasturbine engine components, so that less flow is needed for cooling andefficiency is maintained or increased.

FIG. 3A is a cross-sectional view of gas turbine engine component(turbine or turbomachinery component) 100 with gas path wall 102, takenin a longitudinal direction and that carries a cool first surface 106and an opposite, hot, second surface 108. Cooling hole 104 extendsthrough gas path wall 102 from first surface 106 to second surface 108form cooling hole 60 in the, for example outer wall of an airfoil,casing, combustor liner, exhaust nozzle or other gas turbine enginecomponent, as described above.

Gas path wall 102 of component 100 is exposed to cooling fluid on firstsurface 106 with longitudinal hot gas or working fluid flow H alongsecond surface 108. In some components, for example airfoils, firstsurface 106 is an inner surface (or inner wall) and second surface 108is an outer surface (or outer wall). In other components, for examplecombustor liners and exhaust nozzles, first surface 106 is an outersurface (or outer wall), and second surface 108 is an inner surface (orinner wall). More generally, the terms inner and outer are merelyrepresentative, and may be interchanged.

Cooling hole 104 delivers cooling fluid C from first surface 106 of wall102 to second surface 108, for example to provide diffusive flow andfilm cooling. Cooling hole 104 is also inclined along axis A in adownstream direction, in order to improve cooling fluid coverage oversecond surface 108, with less separation and reduced flow mixing.Longitudinal ridge 124 is provided to reduce flow swirl and flow vortexat outlet 116. Outlet 116 defines a perimeter of cooling hole 104 at anintersection of cooling hole 104 and second surface 108. Surfaces 120,122, 130, and 132 of cooling hole 104 define cooling hole 104 betweeninlet 114 and outlet 116.

As shown in FIG. 3A, cooling hole 104 extends along axis A from inlet114 at first surface 106 of gas path wall 102 to outlet 116 at secondsurface 108. In one embodiment, cooling hole 104 is continuouslydivergent throughout, with continuously increasing cross section or flowarea from inlet 114 through cooling hole 104 to outlet 116. Thus,cooling hole 104 has substantially no convergent or constant-areametering portion between inlet 114 and outlet 116, nor any transitionbetween such a convergent or constant-area metering portion and adiffusion portion. Instead, regulation of cooling flow C is provided byinlet 114, or other geometrical feature at first surface 106, and flowis diffusive through cooling hole 104 over substantially the entirelength from inlet 114 at first surface 106 to outlet 116 at secondsurface 108.

That is, cooling hole 104 is substantially diffusive (or divergent)between inlet 114 and outlet 116, and from first surface 106 to secondsurface 108 of gas path wall 102. As shown in FIG. 3A, for example,upstream and downstream surfaces 120 and 122 of cooling hole 104 divergecontinuously along axis A between inlet 114 and outlet 116. Inparticular, upstream surface 120 and downstream surface 122 diverge awayfrom one another in the longitudinal direction, as defined along hot gasflow H. This increases the cross sectional area (or flow area) ofcooling hole 104, providing diffusive flow to increase the coverage ofcooling fluid C along second surface 108 of gas path wall 102. Inanother embodiment, upstream surface 120 and downstream surface 122cooling hole 104 can converge in the longitudinal direction, and lateralsurfaces 130 and 132 (shown in FIGS. 4A, 4B, 5A, 5B, 6A, 6B, 7A, and 7B)of cooling hole 104 can diverge in the lateral direction.

Longitudinal ridge 124 is formed as a ridge or rib structure alongdownstream surface 122 of cooling hole 104. In this particularconfiguration shown in FIG. 3A, longitudinal ridge 124 extends frominlet 114 to outlet 116, in order to reduce swirl components oversubstantially the entire length of cooling hole 104. Alternatively,longitudinal ridge 124 is truncated between inlet 114 and outlet 116(shown in FIG. 3B), or longitudinal ridge 124 extends along cooling hole104 to form a cusp at inlet 114 on inner surface 106 (shown in FIG. 3C),as described in more detail below.

Longitudinal ridge 124 projects laterally outward from downstreamsurface 122 toward axis A, separating cooling hole 104 into lobes todiscourage swirl flow and reduce flow mixing at outlet 116. Longitudinalridge 124 may also include transition region 128, extending from ridgetransition 118 to trailing edge 126 of outlet 116, as described below(see, e.g., FIGS. 4B, 5A, 5B, 6A, 6B, 7A, 7B). Ridge transition 118 is alocation where longitudinal ridge 124 meets transition region 128. Inthe embodiment illustrated in FIG. 3A, ridge transition 118 has a curvedcross-sectional profile. In alternative embodiments (such as FIGS. 3Aand 3B), ridge transition 118 can have a pointed cross-sectionalprofile. Transition region 128 may be flat/planar, or convex toencourage flow attachment and reduce flow separation or mixing alongsecond surface 108 of gas path wall 102.

FIG. 3B is a cross-sectional view of gas turbine engine component 100with gas path wall 102, showing cooling hole 104 with longitudinal ridge124 in a truncated configuration. In this configuration, longitudinalridge 124 extends from ridge terminus 119 to ridge transition 118. Ridgeterminus 119 is a point where longitudinal ridge 124 meets downstreamsurface 122. Ridge terminus 119 is spaced between inlet 114 and outlet116. This truncated configuration provides diffusive flow from inlet 114through ridge terminus 119 to outlet 116, discourages swirl flow fromridge terminus 119 through outlet 116, and minimizes flow mixing at thesecond surface 108 downstream of outlet 116.

FIG. 3C is a cross-sectional view of gas turbine engine component 100with gas path wall 102, showing cooling hole 104 with a cuspedconfiguration at inlet 114. In this configuration, longitudinal ridge124 projects laterally outward (toward axis A) from downstream surface122 at first surface 106 of gas path wall 102, forming a cusp 125 oninlet 11. In this particular configuration, cusp 125 extends congruentlywith longitudinal ridge 124 along cooling hole 104, from inlet 114toward outlet 116.

The cross-sectional geometry of cooling hole 104 also varies, asdescribed above, and as shown in the figures. The design of cooling hole104 is not limited to these particular examples, however, but alsoencompasses different combinations and variations of the features thatare described, including different features for longitudinal ridge 124,ridge transition 118, ridge terminus 119, and transition region 128.

FIG. 4A is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating a three-lobe geometry for cooling hole 104.Cooling hole 104 extends from inlet 114 on first surface 106 (dashedline) of gas path wall 102 to outlet 116 on second surface 108 (solidline). Second surface 108 is exposed to hot gas flow H in a downstreamlongitudinal direction, as described above.

Lateral (side) surfaces 130 and 132 of cooling hole 104 divergecontinuously along axis A, from inlet 114 at first surface 106 to outlet116 at second surface 108. In particular, side surfaces 130 and 132diverge in a lateral direction, transverse or perpendicular to hot gasflow H, increasing the cross sectional flow area of cooling hole 104 toprovide diffusive flow along substantially the entire passage lengthbetween inlet 114 and outlet 116.

Longitudinal ridges 124 separate cooling hole 104 into lobes 134. Forexample, two ridges 124 may extend along downstream surface 122 todivide cooling hole 104 into three lobes 134, as shown in FIG. 4A. Lobes134 are surfaces of wall 102 which define distinct channel-like portionsof the void of cooling hole 104. Longitudinal ridges 124 project out(upward) from downstream surface 122 toward axis A to discouragetransverse flow components and swirl, reducing flow separation andminimizing flow mixing at the second surface 108 downstream of outlet116.

The geometry of outlet 116 is also selected to improve coolingperformance, including the geometry of trailing edge 126. In particular,outlet 116 may be formed as a delta with arcuate upstream surface 120and substantially straight trailing edge 126, transverse to hot gas flowH. Alternatively, the delta may be configured with a more or less convextrailing edge 126, having a central portion extending downstream alongsecond surface 108 (see FIG. 4B). These configurations further reduceflow separation and increase attachment and laminar flow, for improvedcoverage and cooling efficiency along second surface 108 of gas pathwall 102.

FIG. 4B is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating a two-lobe geometry for cooling hole 104. Inthis configuration, a single longitudinal ridge 124 divides cooling hole104 into two lobes 134, and outlet 116 has a delta configuration witharcuate upstream surface 120 extending toward convex trailing edge 126.

As shown in FIG. 4B, longitudinal ridge 124 extends from inlet 114 toridge transition 118, and transition region 128 extends from ridgetransition 118 to trailing edge 126 of outlet 116. For example,longitudinal ridge 124 may be formed at the intersection or interfacebetween adjacent lobes 134, where lobes 134 have arcuate or curvedsurfaces along downstream surface 122, meeting at a cusped or convexlongitudinal ridge 124. Alternatively, longitudinal ridge 124 may beformed at the intersection or interface between adjacent lobes 134 withsubstantially planar surfaces along downstream surface 122, meeting at atriangular longitudinal ridge 124.

Transition region 128 is defined between arcuate extensions 136 oflongitudinal ridge 124. In the particular configuration of FIG. 4B, forexample, two arcuate extensions 136 form at ridge transition 118 oflongitudinal ridge 124, extending longitudinally and transversely fromridge transition 118 to trailing edge 126 of outlet 116. Transitionregion 128 can be flat or planar. Alternatively, transition region 128can be non-flat and non-planar, such as curved (e.g. convex)longitudinally and/or laterally.

Transition region 128 extends transversely along substantially theentire length of trailing edge 126, between arcuate extensions 136.Alternatively, two or more transition regions 128 extend along trailingedge 126, as defined between three or more lobes 134 (see, e.g., FIG.5A), or a trapezoidal region is provided (FIG. 7B). In each of theseconfigurations, one or more transition regions 128 extend alongsubstantially all of trailing edge 126, eliminating cusps and otherirregularities along trailing edge 126 to encourage attachment andreduce separation for more uniform coverage and higher coolingefficiency. Transition region 128 can further encourage attachment andreduce separation when transition region 128 is convex.

FIG. 5A is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating a truncated lobe geometry for cooling hole104. In this configuration, two longitudinal ridges 124 separate coolinghole 104 into three lobes 134 in the region of outlet 116. Lobes 134extend along cooling hole 104 toward inlet 114, as described above withrespect to FIG. 4A.

As shown in FIG. 5A, however, longitudinal ridges 124 merge at ridgenexus 140, between (circular, oval or elliptical) inlet 114 and(delta-shaped) outlet 116. A single longitudinal ridge 124 then extendsfrom ridge nexus 140 to inlet 114, dividing cooling hole 104 into twolobes in this region.

Considered in the direction of cooling fluid flow, a single longitudinalridge 124 separates cooling hole 104 into two lobes 134 in the regionfrom inlet 114 to ridge nexus 140. Longitudinal ridge 124 splits orbifurcates at ridge nexus 140, dividing cooling hole 104 into threelobes 134 in the region from ridge nexus 140 to outlet 116. Thus, twolongitudinal ridges 124 extend from ridge nexus 140 to two transitions118, with two transition regions 128 extending from transitions 118 totrailing edge 126 of cooling hole 104. Transition regions 128 aredefined between adjacent lobes 134 by arcuate extensions 136, asdescribed above.

FIG. 5B is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating an alternate truncated lobe geometry forcooling hole 104. In this configuration, both longitudinal ridges 124terminate at ridge nexus 140, and no longitudinal ridge 124 extendsbetween inlet 114 and ridge nexus 140. Thus, cooling hole 104 has asingle-lobe configuration from inlet 114 to ridge nexus 140, and athree-lobe configuration from ridge nexus 140 to outlet 116.

FIG. 6A is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating a truncated two-lobe geometry for coolinghole 104. In this configuration, longitudinal ridge 124 is not presentbetween (circular) inlet 114 and ridge terminus 119. Instead,longitudinal ridge 124 separates cooling hole 104 into two lobes 134 inthe region of outlet 116, extending from ridge terminus 119 to ridgetransition 118. Arcuate extensions 136 define a single transition region128, extending between adjacent lobes 134 from ridge transition 118 totrailing edge 126 of outlet 116.

FIG. 6B is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating a cusped inlet geometry for cooling hole104. Cusp 125 discourages swirl and vortex formation, and may furtherrestrict the area of inlet 114 to provide additional flow metering.

In this particular configuration, a single longitudinal ridge 124separates cooling hole 104 into two lobes 134, extending from ridgetransition 118 to inlet 114 so as to define cusp 125 on inlet 114, atfirst surface 106 of gas path wall 102. Cusp 125 thus extendscongruently with longitudinal ridge 124, from inlet 114 to ridgetransition 118 along downstream surface 122 of cooling hole 104. Arcuateextensions 136 extend from ridge transition 118 to trailing edge 126 ofcooling hole 104, defining transition region 128 between adjacent lobes134.

Like ridge 124, cusp 125 projects laterally away from downstream surface122 toward the axis of cooling hole 104 (see FIG. 3C), discouragingtransverse flow components to reduce swirl along cooling hole 104, withless flow mixing at outlet 116 and second surface 108 downstream. Cusp125 also restricts the area of inlet 114, reducing flow vortex andimproving coverage as described above. In continuously diverging designsof cooling hole 104, where flow is diffusive through substantially theentire length of cooling hole 104, cusp 125 provides additional controlof flow metering, where metering is determined primarily by the size andgeometry of inlet 114.

The geometries of longitudinal ridge 124 and cusp 125 vary, and thestructures may be formed as extensions of one another, or distinct. Forexample, one or both of ridge 124 and cusp 125 may be formed as long,narrow features extending along the wall of cooling hole 104 where twosloping sides of lobes 134 meet, or as a narrow raised band or ribstructure between adjacent lobes 134. Ridge 124 and cusp 125 may also beeither substantially pointed or rounded where two curved lobes 134 orwall surfaces meet, or where the direction of curvature reverses along awall of cooling hole 104. Ridge 124 and cusp 125 may also be formed asarched or cone-shape features extending along the boundary of two lobes134.

FIG. 7A is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating an alternate cusped inlet geometry forcooling hole 104. In this configuration, two longitudinal ridges 124extend from cusp 125 at inlet 114 to ridge transition 118, dividingcooling hole 104 into three lobes 134 along substantially the entirepassage length.

FIG. 7B is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating a buried divider or buried ridgeconfiguration for cooling hole 104. In this design, longitudinal ridges124 extend from inlet 114, dividing cooling hole 104 into three lobes134 between first surface 106 of gas path wall and ridge transition 118.As shown in FIG. 7B, however, middle lobe 134 terminates at transitionregion 128, where transition region 128 is bounded between intersections142 with adjacent outer lobes 134.

Unlike arcuate extensions 136 of longitudinal ridges 124, intersections142 do not extend above downstream surface 122 toward axis A of coolinghole 104. Instead, transition region 128 is congruent with downstreamsurface 122, and adjacent lobes 134 curve up from intersections 142toward second (upper) surface 108 of gas path wall 102.

FIG. 7C is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating an alternate cusped inlet geometry forcooling hole 104. In this configuration, two longitudinal ridges 124extend from two cusps 125 at inlet 114 to ridge transition 118, dividingcooling hole 104 into three lobes 134 along substantially the entirepassage length. Thus, inlet 114 has a two-cusped configuration.

FIG. 7C also illustrates a single transition region 128 extending fromlongitudinal ridges 124 to trailing edge 126 of outlet 116. The boundaryof transition region 128 and adjacent lobes 134 are defined alongarcuate extensions 136. Transition region 128 extends acrosssubstantially all of trailing edge 126, eliminating irregularities toprovide more uniform flow coverage and better cooling performance alongsecond surface 108 of gas path wall 102, downstream of outlet 116.Transition region 128 separates lobes 134 from trailing edge 126 ofoutlet 116.

The overall geometry of cooling hole 104 thus varies, as describedabove, and as shown in the figures. The design of inlet 114 and outlet116 may also vary, including various circular, oblate, oval,trapezoidal, triangular, cusped and delta shaped profiles with arcuateor piecewise linear upstream surfaces 120 extending toward straight orconvex trailing edges 126. The configuration of cooling hole 104 is notlimited to these particular examples, moreover, but also encompassesdifferent combinations of the various features that are shown, includinga variety of different cusps 125; transitions 118 with differentcircular, elliptical, oblong and cusped cross sections; and one, two orthree lobes 134, in combination with different transition regions 128bordered by various arcuate extensions 136 and intersections 142.

FIG. 8 is a block diagram illustrating method 200 for forming a coolinghole through the flow path wall of a gas turbine engine component. Forexample, method 200 may be used to form cooling hole 60 or cooling hole104 in the gas path wall of an airfoil, casing, liner, combustor,augmentor or turbine exhaust component of a gas turbine engine or otherturbomachine, as described above.

Method 200 includes forming a cooling hole in a flow path wall of thecomponent (step 202), for example by forming an inlet in the first(cool) surface of the wall (step 204), forming an outlet in the second(hot) surface of the wall (step 206), and forming lobes (step 208)between the inlet and the outlet.

The cooling hole extends along an axis from the inlet to the outlet,diverging continuously along the axis from the first surface to thesecond surface of the flow path wall. The cross-sectional or flow areaof the cooling hole increases continuously from the inlet to the outlet,and flow is diffusive along substantially the entire length of thecooling hole, from the inlet through the cooling hole to the outlet.

Forming a ridge (step 208) comprises forming a longitudinal ridge alonga downstream wall of the cooling hole. The longitudinal ridge dividesthe cooling hole into lobes, for example in the outlet region of thecooling hole. Where the longitudinal ridge extends to the inlet, it maybe formed congruently with a cusp.

In some designs, forming the inlet (step 204) includes forming a cusp inthe inlet (step 210), and in other designs a cusp may be formed alongwith the lobes (step 208). In additional designs, forming the outlet(step 206) may include forming a transition region (step 212). Where atransition region is formed, one or more of the lobes may terminatethere.

The gas turbine engine components, gas path walls and cooling holesdescribed herein can thus be manufactured using one or more of a varietyof different processes. These techniques provide each cooling hole withits own particular configuration and features, including, but notlimited to, inlet, metering, transition, diffusion, outlet, upstreamsurface, downstream surface, lateral surface, longitudinal, lobe anddownstream edge features, as described above. In some cases, multipletechniques can be combined to improve overall cooling performance orreproducibility, or to reduce manufacturing costs.

Suitable manufacturing techniques for forming the cooling configurationsdescribed here include, but are not limited to, electrical dischargemachining (EDM), laser drilling, laser machining, electrical chemicalmachining (ECM), water jet machining, casting, conventional machiningand combinations thereof. Electrical discharge machining includes bothmachining using a shaped electrode as well as multiple pass methodsusing a hollow spindle or similar electrode component. Laser machiningmethods include, but are not limited to, material removal by ablation,trepanning and percussion laser machining. Conventional machiningmethods include, but are not limited to, milling, drilling and grinding.

The gas flow path walls and outer surfaces of some gas turbine enginecomponents include one or more coatings, such as bond coats, thermalbarrier coatings, abrasive coatings, abradable coatings and erosion orerosion-resistant coatings. For components having a coating, the inlet,transition, and outlet cooling features may be formed prior to a coatingapplication, after a first coating (e.g., a bond coat) is applied, orafter a second or third (e.g., interlayer) coating process, or a finalcoating (e.g., environmental or thermal barrier) process. Depending oncomponent type, cooling hole or passage location, repair requirementsand other considerations, the outlet features may be located within awall or substrate, within a thermal barrier coating or other coatinglayer applied to a wall or substrate, or combinations thereof. Thecooling geometry and other features may remain as described above,regardless of position relative to the wall and coating materials orairfoil materials.

While the invention is described with reference to exemplaryembodiments, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted withoutdeparting from the spirit and scope of the invention. In addition,different modifications may be made to adapt the teachings of theinvention to particular situations or materials, without departing fromthe essential scope thereof. The invention is thus not limited to theparticular examples disclosed herein, but includes all embodimentsfalling within the scope of the appended claims.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments ofthe present invention.

A component for a gas turbine engine can include a gas path wall havinga first surface, a second surface exposed to hot gas flow, and a coolinghole extending through the gas path wall. The cooling hole can includean inlet formed in the first surface, an outlet formed in the secondsurface, cooling hole surfaces that define the cooling hole between theinlet and the outlet, and a longitudinal ridge formed along at least oneof the cooling hole surfaces. The longitudinal ridge can separate thecooling hole into first and second lobes. The cooling hole can divergethrough the gas path wall, such that cross-sectional area of the coolinghole increases continuously from the inlet through the cooling hole tothe outlet.

The component of the preceding paragraph can optionally include,additionally and/or alternatively any, one or more of the followingfeatures, configurations and/or additional components:

the longitudinal ridge can extend from the inlet along at least one ofthe cooling hole surfaces toward the outlet, dividing the cooling holeinto the first and second lobes in a region of the inlet;

a cusp can be on the inlet and the longitudinal ridge can extend fromthe cusp;

a ridge terminus can be spaced along one of the cooling hole surfacesbetween the inlet and the outlet, the longitudinal ridge can terminateat the ridge terminus, and the longitudinal ridge can extend from theridge terminus toward the outlet;

a ridge transition can be spaced along the longitudinal ridge betweenthe inlet and the outlet, and a transition region of the cooling holecan extend from the ridge transition to a trailing edge of the outlet;

the longitudinal ridge can be a first longitudinal ridge such that firstand second longitudinal ridges can divide the cooling hole into first,second, and third lobes;

a ridge nexus can be spaced along the cooling hole between the inlet andthe outlet, the first and second longitudinal ridges can meet at theridge nexus, the second lobe can be positioned between the first andthird lobes, and the second lobe can truncate at the ridge nexus;

the first and second ridges can join at the ridge nexus to form a singlelongitudinal ridge that extends from the ridge nexus toward the inlet;

the first and second longitudinal ridges can terminate at the ridgenexus such that no longitudinal ridge extends from the ridge nexustoward the inlet; and/or

the second surface can form one of a pressure surface, a suction surfaceor a platform surface of an airfoil.

An airfoil can include a flow path wall having a first surface exposedto cooling fluid and a second surface exposed to hot gas flow. A coolinghole can be formed in the flow path wall and be laterally divergingcontinuously from an inlet at the first surface to an outlet at thesecond surface. A longitudinal ridge can be defined along the coolinghole and divide the cooling hole into first and second lobes. Flow ofthe cooling fluid can be substantially diffusive through the coolinghole, from the inlet at the first surface of the flow path wall throughto the outlet at the second surface of the flow path wall.

The airfoil of the preceding paragraph can optionally include,additionally and/or alternatively any, one or more of the followingfeatures, configurations and/or additional components:

the cooling hole can diverge continuously in longitudinal and transversedirections from the inlet to the outlet, and the longitudinal andtransverse directions can be defined with respect to the hot gas flow;

the cooling hole can be inclined in a downstream direction between thefirst surface and the second surface;

the outlet can have a substantially straight or convex trailing edge;

the longitudinal ridge can be a first longitudinal ridge such that firstand second longitudinal ridges can divide the cooling hole into first,second, and third lobes; and/or

a gas turbine engine can include the airfoil.

A gas turbine engine component can include a gas path wall, a coolinghole extending through the gas path wall, and a longitudinal ridgeextending along the cooling hole. The gas path wall can have a firstsurface and a second surface exposed to hot gas flow. The cooling holecan be continuously diverging from an inlet in the first surface to anoutlet in the second surface. The longitudinal ridge can divide thecooling hole into first and second lobes.

The gas turbine engine component of the preceding paragraph canoptionally include, additionally and/or alternatively any, one or moreof the following features, configurations and/or additional components:

a cusp can be formed on the inlet, and the longitudinal ridge can extendfrom the cusp toward the outlet along a downstream wall of the coolinghole;

the longitudinal ridge can be a first longitudinal ridge such that firstand second longitudinal ridges can divide the cooling hole into first,second, and third lobes; and/or

a ridge transition can be spaced along the longitudinal ridge betweenthe inlet and the outlet, and a planar or convex transition region canextend along the cooling hole from the ridge transition to a trailingedge of the outlet.

1. A component for a gas turbine engine, the component comprising: a gaspath wall having a first surface and a second surface, wherein thesecond surface is exposed to hot gas flow; and a cooling hole extendingthrough the gas path wall, the cooling hole comprising: an inlet formedin the first surface; an outlet formed in the second surface; coolinghole surfaces that define the cooling hole between the inlet and theoutlet; and a longitudinal ridge formed along at least one of thecooling hole surfaces, the longitudinal ridge separating the coolinghole into first and second lobes; wherein the cooling hole divergesthrough the gas path wall, such that cross-sectional area of the coolinghole increases continuously from the inlet through the cooling hole tothe outlet.
 2. The component of claim 1, wherein the longitudinal ridgeextends from the inlet along at least one of the cooling hole surfacestoward the outlet, dividing the cooling hole into the first and secondlobes in a region of the inlet.
 3. The component of claim 2, and furthercomprising: a cusp on the inlet, wherein the longitudinal ridge extendsfrom the cusp.
 4. The component of claim 1, and further comprising: aridge terminus spaced along one of the cooling hole surfaces between theinlet and the outlet, wherein the longitudinal ridge terminates at theridge terminus, and wherein the longitudinal ridge extends from theridge terminus toward the outlet.
 5. The component of claim 1, andfurther comprising: a ridge transition spaced along the longitudinalridge between the inlet and the outlet; and a transition region of thecooling hole extending from the ridge transition to a trailing edge ofthe outlet.
 6. The component of claim 1, wherein the longitudinal ridgeis a first longitudinal ridge, and further comprising: a secondlongitudinal ridge; and a third lobe, wherein the first and secondlongitudinal ridges divide the cooling hole into the first, second, andthird lobes.
 7. The component of claim 6, and further comprising: aridge nexus spaced along the cooling hole between the inlet and theoutlet, wherein the first and second longitudinal ridges meet at theridge nexus, wherein the second lobe is positioned between the first andthird lobes, and wherein the second lobe truncates at the ridge nexus.8. The component of claim 7, wherein the first and second ridges join atthe ridge nexus to form a single longitudinal ridge that extends fromthe ridge nexus toward the inlet.
 9. The component of claim 6, whereinthe first and second longitudinal ridges terminate at the ridge nexussuch that no longitudinal ridge extends from the ridge nexus toward theinlet.
 10. The component of claim 1, wherein the second surface formsone of a pressure surface, a suction surface or a platform surface of anairfoil.
 11. An airfoil comprising: a flow path wall having a firstsurface exposed to cooling fluid and a second surface exposed to hot gasflow; a cooling hole formed in the flow path wall, the cooling holelaterally diverging continuously from an inlet at the first surface toan outlet at the second surface; and a longitudinal ridge defined alongthe cooling hole, the longitudinal ridge structure dividing the coolinghole into first and second lobes; wherein flow of the cooling fluid issubstantially diffusive through the cooling hole, from the inlet at thefirst surface of the flow path wall through to the outlet at the secondsurface of the flow path wall.
 12. The airfoil of claim 11, wherein thecooling hole diverges continuously in longitudinal and transversedirections from the inlet to the outlet, the longitudinal and transversedirections defined with respect to the hot gas flow.
 13. The airfoil ofclaim 11, wherein the cooling hole is inclined in a downstream directionbetween the first surface and the second surface.
 14. The airfoil ofclaim 11, wherein the outlet has a substantially straight or convextrailing edge.
 15. The airfoil of claim 11, wherein the longitudinalridge is a first longitudinal ridge, and further comprising: a secondlongitudinal ridge; and a third lobe, wherein the first and secondlongitudinal ridges divide the cooling hole into the first, second, andthird lobes.
 16. A gas turbine engine comprising the airfoil of claim11.
 17. A gas turbine engine component comprising: a gas path wallhaving a first surface and a second surface, the second surface exposedto hot gas flow; a cooling hole extending through the gas path wall, thecooling hole continuously diverging from an inlet in the first surfaceto an outlet in the second surface; and a longitudinal ridge extendingalong the cooling hole, the longitudinal ridge dividing the cooling holeinto first and second lobes.
 18. The gas turbine engine component ofclaim 17, further comprising: a cusp formed on the inlet, wherein thelongitudinal ridge extends from the cusp toward the outlet along adownstream wall of the cooling hole.
 19. The gas turbine enginecomponent of claim 17, wherein the longitudinal ridge is a firstlongitudinal ridge, and further comprising: a second longitudinal ridge;and a third lobe, wherein the first and second longitudinal ridgesdivide the cooling hole into the first, second, and third lobes
 20. Thegas turbine engine component of claim 17, further comprising: a ridgetransition spaced along the longitudinal ridge between the inlet and theoutlet; and a planar or convex transition region extending along thecooling hole from the ridge transition to a trailing edge of the outlet.